It is known practice for selected gas turbine engine components, especially in the turbine section, to be internally air cooled by a supply of air bleed from a compressor offtake. Such cooling is necessary to maintain component temperatures within the working range of the materials from which they are constructed. Higher engine gas temperature have led to increased cooling bleed requirements resulting in reduced cycle efficiency and increased emission levels.
To date, it has been possible to improve the design of cooling systems to minimize cooling flow at relative low cost. In the future, engine temperatures will increase to levels at which it is necessary to have complex cooling features to maintain low cooling flows.
An effective cooling system of blades for a gas turbine engine is disclosed in U.S. Pat. No. 5,720,431. The disclosed airfoil includes a double wall configuration in the mid-chord region with a plurality of radial feed passages defined on each side of the airfoil between the inner wall and the outer wall. A central radially extending feed chamber is defined between the two inner walls. A trailing edge of the airfoil includes a conventional single wall configuration with two outer walls defining a sequence of trailing edge cavities extending radially through the airfoil and being axially connected fluidly such that a common exhaust port discharge at the trailing edge directly. Due to the bent airfoil profile there is a large material accumulation at the end of the pressure side cavity which leads to a higher temperature gradient in the airfoil.
The same negative aspect of material accumulation at the pressure sided trailing edge region of an airfoil can be observed at the known air cooled airfoil disclosed in EP 1 267 038 B1. The herein described airfoil provides an axially orientated suction sided near wall channel which discharges its cooling air at the trailing edge towards the pressure side. As the trailing edge is subject to a very high heat load, the suction side cooling channel has to provide sufficient air to keep the trailing edge temperatures sufficiently low.
Another design for internally cooling an airfoil for gas turbine engine is disclosed in U.S. Pat. No. 7,946,815 B2 which provides near wall cooling channels to keep the wall temperatures low enough to provide sufficient component life. Separate channels at the pressure side and the suction side are for cooling the outer side of the airfoil which is exposed to the hot gas flow in a gas turbine stage. The known airfoil disclosed in the before document comprises a suction and a pressure side wall each extending in an axial direction, which means from a leading to a trailing edge region of the airfoil. The known airfoil further comprises a suction wall sided cooling channel extending in axial direction confined by a suction side wall and a first inner wall, as well a pressure wall sided cooling channel extending in axial direction confined by the pressure side wall and a second inner wall. The first and second inner wall borders some feed chambers, some of them are fluidly connected, for feeding said at least one suction and pressure sided cooling channel with a cooling medium, preferably compressed air each by a multitude of through holes inside of said first and second inner wall.